This invention relates to axial flow gas turbine engine compressors and fans and, more particularly, to compressor stages in which the relative flow velocity entering the rotating blade row is supersonic.
The purpose of the compressor or fan in a gas turbine engine is to raise the pressure and reduce the volume of air as it is pumped through the engine. The compressor comprises a plurality of axially stacked stages, each stage consisting of a row of rotating blades (rotor blades) followed by a row of stationary blades (stators). Within each stage the airflow is accelerated through the rotor and decelerated through the stator with a resulting pressure rise, the pressure ratio being multiplied by each succeeding stage. The cross-sectional area of the compressor decreases gradually through the compressor from the low to the high pressure end in order to maintain the axial velocity of the air as pressure increases. Each rotor blade and stator are of airfoil section, and the number and size of each change from stage to stage as the air passages through the compressor are gradually diminished.
The work input of the rotor (i.e., the work done on the air) is generally a function of the difference of the square of the absolute velocities exiting and entering the rotor, plus the difference of the square of the relative velocities entering and exiting the rotor. Thus, the energy input per stage can increase by decreasing the inlet absolute velocity and the discharge relative velocity, or by increasing the inlet relative velocity and discharge absolute velocity. Clearly, both alternatives are limited. Inlet absolute velocity is determined by the flow capacity of the compressor and cannot be easily adjusted in a practical manner to control work input. The possiblity of decreasing the discharge relative velocity is limited by the permissible adverse pressure gradient in the rotor passages between blades. Therefore, the attainment of higher compression ratios per stage is more readily attained by increasing the values of inlet relative velocities and discharge absolute velocity. Since these velocities are, in essence, the relative velocities into both the rotor and the following stator passages, large shock losses may be produced if these velocities are permitted to become supersonic. On the other hand, the requirement that the rotor inlet relative velocity be subsonic imposes limitations on the peripheral speed of the rotor. Thus, the requirements of high compressor capacity and high stage compression ratio cannot be satisfied in a conventional subsonic machine. It is clear, then, that supersonic compressors (i.e., axial flow compressors having supersonic relative velocities at the rotor blade leading edge) offer considerable savings in weight and size for any predetermined flow capacity and pressure ratio, but only provided that they are also so designed as to reduce the severity of the shock losses.